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At least 847 records · Page 47

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NASA NTRS1992

Propulsion systems using in situ propellants for a Mars Ascent Vehicle

The indigenous propellants of oxygen and carbon monoxide were studied for use in a Mars Ascent Vehicle (MAV). Both the oxygen and carbon monoxide were evaluated as turbine working fluids for a full expander engine cycle used in the MAV. Two oxygen working fluid engines and four carbon monoxide engines were investigated. The maximum hot-gas-side wall temperature was limited to either 445 K (800R) or 556 K (1000 R) over a range of working fluid mass flows. An engine thrust of 44.5 kN (1,000 lb(sub f)) was assumed. The engine characteristics of coolant inlet pressure, total engine length, specific impulse, pump efficiencies, and turbine power required were optimized. Of the six engines evaluated, the study showed that either type of working fluid was feasible for a MAV engine. Using carbon monoxide as the working fluid resulted in the best engine characteristics.

Wadel, Mary F.↗

NASA NTRS1992

Propulsion system requirements for reusable single-stage-to-orbit rocket vehicles

The conceptual design of a single-stage-to-orbit (SSTO) vehicle using a wide variety of evolutionary technologies has recently been completed as a part of NASA's Advanced Manned Launch System (AMLS) study. The employment of new propulsion system technologies is critical to the design of a reasonably sized, operationally efficient SSTO vehicle. This paper presents the propulsion system requirements identified for this near-term AMLS SSTO vehicle. Sensitivities of the vehicle to changes in specific impulse and sea-level thrust-to-weight ratio are examined. The results of a variety of vehicle/propulsion system trades performed on the near-term AMLS SSTO vehicle are also presented.

Stanley, Douglas O.↗

NASA NTRS1992

Test experience, 490 N high performance (321 sec Isp) engine

Engines with area ratios of 44:1 and 286:1 are tested by means of hot fire tests using the NTO/MMH bipropellant to maximize the performance of the combined technologies. The low-thrust engine systems are designed with oxidation resistant materials that can operate at temperatures of more than 2204 C for tens of hours. The chamber is attached to the injector in a configuration that prevents overheating of the injector, valve, and the spacecraft interface. Three injectors with 44:1 area ratios are capable of nominal specific impulse values of 309 sec, and a performance of 321 lbf-sec/lbm is noted for an all-welded engine assembly with area ratio of 286:1. The all-welded engine is shown to have an acceptable design margin for thermal characteristics. High-performance liquid apogee engines are shown to perform optimally when based on iridium/rhenium chamber technology, use of a special platelet injector, and the minimization of losses due to fuel-film cooling.

Schoenman, L.↗

NASA NTRS1991

Nuclear rocket propulsion: NASA plans and progress - FY 1991

NASA has initiated planning for a technology development project for nuclear rocket propulsion systems for space exploration initiative (SEI) human and robotic missions to the Moon and to Mars. An interagency project is underway that includes the Department of Energy National Laboratories for nuclear technology development. The activities of the project planning team in FY 1990 and 1991 are summarized. The progress to date is discussed, and the project plan is reviewed. Critical technology issues were identified and include: (1) nuclear fuel temperature, life, and reliability; (2) nuclear system ground test; (3) safety; (4) autonomous system operation and health monitoring; and (5) minimum mass and high specific impulse.

Clark, John S.↗

NASA NTRS1991

In-situ carbon dioxide fixation on Mars

The authors examine several novel proposals for CO2 fixation through chemical, photochemical, and photoelectrochemical means. For example, the reduction of CO2 to hydrocarbons such as acetylene (C2H2) can be accomplished with hydrogen. Acetylene has a theoretical vacuum specific impulse of about 375 s. The authors also examine potential uses of CO, as obtained or further reduced to carbon, as a reducing agent in metal oxide processing to form metals or metal carbides for use as structural or power materials. The CO2 can be recycled to generate O2 and CO. In this study, the authors examine reaction schemes for processing in situ resources. The authors highlight chemistry with hydrogen and carbon monoxide to produce propellants and other necessities of manned exploration.

Hepp, Aloysius F.↗

NASA NTRS1992

Space transfer with ground-based laser/electric propulsion

Ground-based high-power CW lasers can be used to beam power to photovoltaic receivers in space that furnish electricity to space vehicles; this energy can also be used to power electric-propulsion orbital transfer vehicles. An account is presently given of the anticipated requirements for the pulsed FEL lasers, large adaptive optics, photovoltaic receivers, and high specific impulse electrical propulsion. Preliminary system analysis results are presented.

Landis, Geoffrey A.↗

NASA NTRS1992

Dependence of hydrogen arcjet operation on electrode geometry

The dependence of 2kW hydrogen arcjet performance on cathode to anode electrode spacing was evaluated at specific impulses of 900 and 1000 s. Less than 2 absolute percent change in efficiency was measured for the spacings tested which did not repeat the 14 absolute percent variation reported in earlier work with similar electrode designs. A different nozzle configuration was used to quantify the variation in hydrogen arcjet performance over an extended range of electrode spacing. Electrode gap variation resulted in less than 3 absolute percent change in efficiency. These null results suggested that electrode spacing is decoupled from hydrogen arcjet ignition. The dependence of breakdown voltage on mass flow rate and electrode agreed with Paschen curves for hydrogen. Preliminary characterization of the dependence of hydrogen arcjet ignition on rates of pulse repetition and pulse voltage rise were also included for comparison with previous results obtained using simulated hydrazine.

Pencil, Eric J.↗

NASA NTRS1992

Numerical study of low pressure nuclear thermal rockets

The flowfields and performance of low pressure nuclear thermal rockets, which use hydrogen as a propellant, are studied by solving the Navier-Stokes equations and the species equations. A finite-rate chemistry model is used in the species equations, and the turbulence is simulated by the Baldwin-Lomax turbulence model with a modified van Driest's damping constant. The calculated results for the chamber temperatures of 3200 K and 4000 K with a chamber pressure range of 0.1 atm to 6 atm are presented as contours, centerline variations, and exit profiles. The performance values from the present calculations, such as the vacuum specific impulse and thrust, are compared with those from the 1D, inviscid equilibrium and frozen flow code.

Kim, Suk C.↗

NASA NTRS1992

Results of a 1462 hour ammonia arcjet endurance test

A total of 1462 hours of operaton were demonstrated in an endurance test of a 30 kWe-class ammonia arcjet operated at 10 kWe. The propellant flow rate was 0.170 g/s, and the measured performance increased from approximately 650 s specific impulse at 36 percent efficiency at the beginning of the test to 675 s at 39 percent near the end. The voltage increased and the current dropped slightly over the first 400 hours, and then remained approximately constant for the remainder of the test. The test, scheduled for 1500 hours, was terminated after an insulator in the rear of the engine failed. Post-test examination of the electrodes revealed only minimal damage. Although there was evidence of a number of mass transport processes occurring in the discharge chamber, the primary life-limiting wear mechanisms appear to be cathode tip erosion and constrictor melting.

Polk, J. E.↗

NASA NTRS1992

Ammonia arcjet engine behavior in a cyclic endurance test at 10 kW

The behavior of a 30 kWe-class ammonia arcjet operated at 10 kWe during the 707 successful cycles of an endurance test is described. The propellant flow rate was 0.170 g/s, and the measured performance was about 630 s specific impulse at an efficiency of 0.34. Data obtained indicate that the terminal voltage increased over the first 300 cycles, and then remained approximately constant for the remainder of the test, which suggests that the cathode eroded initially and then reached a stable geometry. No major changes were observed in thruster performance. The test was terminated by a series of external arcs.

Polk, J. E.↗

NASA NTRS1992

Low thrust chemical rocket technology

A technology program aimed at improving the performance of low thrust chemical rockets for spacecraft onboard applications is reviewed. Navier-Stokes analyses of low Reynolds number rocket flows have been compared with local flow property measurements obtained using Rayleigh and Raman diagnostics in a 100 N gaseous hydrogen/gaseous oxygen rocket. It is indicated that computational domain should include the near injector flow and that the shear layer combustion model needs improvement. The system analyses and technical efforts intended to develop a technology base for higher performance propellants are presented. A LOX/hydrazine engine is demonstrated to have a 95 percent theoretical c-star which translates into a projected vacuum specific impulse of 345 seconds at an area ratio of 204:1.

Schneider, Steven J.↗

NASA NTRS1993

Nuclear thermal rocket nozzle testing and evaluation program

Performance characteristics of the Nuclear Thermal Rocket can be enhanced through the use of unconventional nozzles as part of the propulsion system. The Nuclear Thermal Rocket nozzle testing and evaluation program being conducted at the NASA Lewis is outlined and the advantages of a plug nozzle are described. A facility description, experimental designs and schematics are given. Results of pretest performance analyses show that high nozzle performance can be attained despite substantial nozzle length reduction through the use of plug nozzles as compared to a convergent-divergent nozzle. Pretest measurement uncertainty analyses indicate that specific impulse values are expected to be within + or - 1.17 pct.

Davidian, Kenneth O.↗

NASA NTRS1992

Combined high and low thrust propulsion for fast piloted Mars missions

The mission benefits of using both high thrust nuclear thermal propulsion (NTP) and low acceleration, high specific impulse nuclear electric propulsion (NEP) to reduce piloted trip times to Mars with reasonable initial mass are assessed. Recent updates in mission design, such as the Earth fly-by return, are assessed for their impact on previous studies. In addition, the Synthesis Commission split mission to Mars in 2014 is also assessed using combined propulsion. Results show an 80 to 100 day reduction in trip time over the reference NTP or NEP systems and missions, with comparable or reduced vehicle initial masses. The impacts of the mission and system analyses upon technology planning and design are discussed.

Gilland, James H.↗

NASA NTRS1992

Electric propulsion

Electric propulsion (EP) is an attractive option for unmanned orbital transfer vehicles (OTV's). Vehicles with solar electric propulsion (SEP) could be used routinely to transport cargo between nodes in Earth, lunar, and Mars orbit. Electric propulsion systems are low-thrust, high-specific-impulse systems with fuel efficiencies 2 to 10 times the efficiencies of systems using chemical propellants. The payoff for this performance can be high, since a principal cost for a space transportation system is that of launching to low Earth orbit (LEO) the propellant required for operations between LEO and other nodes. Several aspects of electric propulsion, including candidate systems and the impact of using nonterrestrial materials, are discussed.

Garrison, Philip W.↗

NASA NTRS1992

Conceptual analysis of a lunar base transportation system

Important to the planning for a lunar base is the development of transportation requirements for the establishment and maintenance of that base. This was accomplished as part of a lunar base systems assessment study conducted by the NASA Langley Research Center in conjunction with the NASA Johnson Space Center. Lunar base parameters are presented using a baseline lunar facility concept and timeline of developmental phases. Masses for habitation and scientific modules, power systems, life support systems, and thermal control systems were generated, assuming space station technology as a starting point. The masses were manifested by grouping various systems into cargo missions and interspersing manned flights consistent with construction and base maintenance timelines. A computer program that sizes the orbital transfer vehicles (OTV's), lunar landers, lunar ascenders, and the manned capsules was developed. This program consists of an interative technique to solve the rocket equation successively for each velocity correction (delta V) in a mission. The delta V values reflect integrated trajectory values and include gravity losses. As the program computed fuel masses, it matched structural masses from General Dynamics' modular space-based OTV design. Variables in the study included the operation mode (i.e., expendable vs. reusable and single-stage vs. two-stage OTV's), cryogenic specific impulse, reflecting different levels of engine technology, and aerobraking vs. all-propulsive return to Earth orbit. The use of lunar-derived oxygen was also examined for its general impact. For each combination of factors, the low-Earth orbit (LEO) stack masses and Earth-to-orbit (ETO) lift requirements are summarized by individual mission and totaled for the developmental phase. In addition to these discrete data, trends in the variation of study parameters are presented.

Hoy, Trevor D.↗

NASA NTRS1992

Electric propulsion for lunar exploration and lunar base development

Using electric propulsion to deliver materials to lunar orbit for the development and construction of a lunar base was investigated. Because the mass of the base and its life-cycle resupply mass are large, high specific impulse propulsion systems may significantly reduce the transportation system mass and cost. Three electric propulsion technologies (arcjet, ion, and magnetoplasmadynamic (MPD) propulsion) were compared with oxygen/hydrogen propulsion for a lunar base development scenario. Detailed estimates of the orbital transfer vehicles' (OTV's) masses and their propellant masses are presented. The fleet sizes for the chemical and electric propulsion systems are estimated. Ion and MPD propulsion systems enable significant launch mass savings over O2/H2 propulsion. Because of the longer trip time required for the low-thrust OTV's, more of them are required to perform the mission model. By offloading the lunar cargo from the manned O2/H2 OTV missions onto the electric propulsion OTV's, a significant reduction of the low Earth orbit (LEO) launch mass is possible over the 19-year base development period.

Palaszewski, Bryan↗

NASA NTRS1992

Lunar He-3, fusion propulsion, and space development

The recent identification of a substantial lunar resource of the fusion energy fuel He-3 may provide the first terrestrial market for a lunar commodity and, therefore, a major impetus to lunar development. The impact of this resource-when burned in D-He-3 fusion reactors for space power and propulsion-may be even more significant as an enabling technology for safe, efficient exploration and development of space. One possible reactor configuration among several options, the tandem mirror, illustrates the potential advantages of fusion propulsion. The most important advantage is the ability to provide either fast, piloted vessels or high-payload-fraction cargo vessels due to a range of specific impulses from 50 sec to 1,000,000 sec at thrust-to-weight ratios from 0.1 to 5x10(exp -5). Fusion power research has made steady, impressive progress. It is plausible, and even probable, that fusion rockets similar to the designs presented here will be available in the early part of the twenty-first century, enabling a major expansion of human presence into the solar system.

Santarius, John F.↗

NASA NTRS1992

Advanced small rocket chambers: Option 1, 14 lbf Ir-Re rocket

A high performance Ir-Re 14 lbf (62 N) chamber and nozzle which can be a direct replacement for a production engine was designed, built, hot fired and vibration acceptance tested. It passed all acceptance tests satisfactorily and demonstrated a 20 sec increase in specific impulse (Is) over the conventional 14 lbf silicide coated Cb chamber. The high performance engine uses the production valve and injector without modification. Incorporation of a secondary mixing device or Boundary Layer Trip within the combustion chamber results in elimination of the fuel film coolant, improvement in flow uniformity, the 20 sec performance increase, and reduction of a potential source of spacecraft contamination. Measured Is was 305 sec at 75:1 area ratio, with monomenthylhydrazine and nitrogen tetroxide propellants. Qualification tests remain to be done.

Jassowski, Donald M.↗